U.S. patent number 10,773,798 [Application Number 15/720,626] was granted by the patent office on 2020-09-15 for rotor hub with blade-to-blade dampers attached to the pitch change axis.
This patent grant is currently assigned to Bell Textron Inc.. The grantee listed for this patent is Bell Helicopter Textron Inc.. Invention is credited to Andrew Paul Haldeman, Dalton T. Hampton, Bryan Marshall, Frank Bradley Stamps.
United States Patent |
10,773,798 |
Haldeman , et al. |
September 15, 2020 |
Rotor hub with blade-to-blade dampers attached to the pitch change
axis
Abstract
An aircraft rotor assembly has a yoke and a plurality of rotor
blade assemblies coupled thereto. Each of the rotor blade
assemblies include a rotor blade, a bearing, and a blade grip
coupling the rotor blade to the bearing. Each of the rotor blades
is rotatable about a lead-lag axis, flap axis, and a pitch change
axis, wherein all the axes intersect within the bearing. Adjacent
pairs of rotor blade assemblies are coupled together via a damper
assembly that is coupled to the pitch change axis of each of the
rotor blade assemblies.
Inventors: |
Haldeman; Andrew Paul (Fort
Worth, TX), Hampton; Dalton T. (Fort Worth, TX),
Marshall; Bryan (Mansfield, TX), Stamps; Frank Bradley
(Colleyville, TX) |
Applicant: |
Name |
City |
State |
Country |
Type |
Bell Helicopter Textron Inc. |
Fort Worth |
TX |
US |
|
|
Assignee: |
Bell Textron Inc. (Fort Worth,
TX)
|
Family
ID: |
65897197 |
Appl.
No.: |
15/720,626 |
Filed: |
September 29, 2017 |
Prior Publication Data
|
|
|
|
Document
Identifier |
Publication Date |
|
US 20190100300 A1 |
Apr 4, 2019 |
|
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
B64C
27/72 (20130101); B64C 27/39 (20130101); B64C
27/48 (20130101); B64C 27/35 (20130101); B64C
27/001 (20130101); B64C 27/635 (20130101); B64C
27/605 (20130101); B64C 2027/7255 (20130101); B64C
27/06 (20130101) |
Current International
Class: |
B64C
27/39 (20060101); B64C 27/00 (20060101); B64C
27/605 (20060101); B64C 27/72 (20060101); B64C
27/48 (20060101); B64C 27/06 (20060101) |
Field of
Search: |
;244/17.25 |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
Primary Examiner: Brown; Claude J
Attorney, Agent or Firm: Lightfoot & Alford PLLC
Claims
What is claimed is:
1. An aircraft rotor assembly, comprising: a yoke configured to be
coupled to and rotate with a mast about a mast axis, the yoke
defining a plurality of bearing pockets; a plurality of rotor blade
assemblies, each rotor blade assembly comprising: a rotor blade; a
bearing, wherein the bearing is at least partially disposed within
one of the plurality of bearing pockets, the bearing forming a
lead-lag hinge with a lead-lag axis that is substantially parallel
to the mast axis, a flap hinge with a flap axis that is
substantially perpendicular to and intersects the lead-lag axis,
and a pitch change hinge with a pitch change axis that is
substantially perpendicular to and intersects both the lead-lag
axis and the flap axis; and a blade grip, the blade grip coupling
the rotor blade to the bearing, the blade grip including a shaft
between the yoke and the rotor blade, wherein the shaft intersects
the pitch change axis; and a plurality of damper assemblies, each
of the plurality of damper assemblies including a first end and a
second end, the first end of each damper assembly and the second
end of an adjacent damper assembly both being connected to a single
concentric rod end bearing, each concentric rod end bearing
comprising a ball and concentric inner and outer races, each race
being connected to either the first end of one of the damper
assemblies or to the second end of an adjacent damper assembly,
each of the concentric rod end bearings being coupled to the shaft
of one of the plurality of blade grips such that a center point of
the ball of each of the concentric rod end bearings is
approximately coincident to the pitch change axis of each of the
plurality of rotor blade assemblies.
2. The aircraft rotor assembly of claim 1, wherein each of the
plurality of damper assemblies do not include another joint between
the concentric rod end bearings to which it is attached.
3. The aircraft rotor assembly of claim 1, wherein the blade grip
includes a central portion connecting an upper plate of the blade
grip to a lower plate of the blade grip, the central portion being
located between the shaft and the rotor blade.
4. The aircraft rotor assembly of claim 1, further comprising: a
control system for collective and cyclic control of a pitch of each
of the plurality of rotor blade assemblies.
5. The aircraft rotor assembly of claim 4, the control system,
comprising: a pitch horn coupled to each of the blade grips; a
swashplate; and a pitch link coupled to each of the pitch horns and
the swashplate, each of the pitch links being located closer to the
mast axis than an outermost surface of the yoke.
6. The aircraft rotor assembly of claim 5, wherein each of the
pitch links is located closer to the mast axis than the outermost
portion of the bearing pockets.
7. The aircraft rotor assembly of claim 6, wherein the yoke is
constructed of a composite material.
8. The aircraft rotor assembly of claim 7, wherein each of the
plurality of rotor blade assemblies are able to rotate about the
lead-lag hinge by at least 1 degree in a lead direction and at
least 1 degree in a lag direction.
9. An aircraft, comprising: a fuselage; a powerplant; a mast
coupled to the powerplant; and a rotor assembly, comprising: a yoke
coupled to the mast and being configured to rotate about a mast
axis, the yoke defining a plurality of bearing pockets; a plurality
of rotor blade assemblies, each rotor blade assembly comprising: a
rotor blade; a bearing, wherein the bearing is at least partially
disposed within one of the plurality of bearing pockets, the
bearing forming a lead-lag hinge with a lead-lag axis that is
substantially parallel to the mast axis, a flap hinge with a flap
axis that is substantially perpendicular to and intersects the
lead-lag axis, and a pitch change hinge with a pitch change axis
that is substantially perpendicular to and intersects both the
lead-lag axis and the flap axis; and a blade grip, the blade grip
coupling the rotor blade to the bearing, the blade grip including a
shaft between the yoke and the rotor blade, wherein the shaft
intersects the pitch change axis; and a plurality of damper
assemblies, each of the plurality of damper assemblies including a
first end and a second end, the first end of each of the plurality
of damper assemblies and the second end of an adjacent damper
assembly both being connected to a single concentric rod end
bearing, each concentric rod end bearing comprising a ball and
concentric inner and outer races, each race being connected to
either the first end of one of the damper assemblies or to the
second end of an adjacent damper assembly, each of the concentric
rod end bearings being coupled to the shaft of one of the plurality
of blade grips such that a center point of the ball of each of the
concentric rod end bearings is approximately coincident to the
pitch change axis of each of the plurality of rotor blade
assemblies.
10. The aircraft of claim 9, further comprising: a control system
for collective and cyclic control of a pitch of each of the
plurality of rotor blade assemblies.
11. The aircraft of claim 10, the control system, comprising: a
pitch horn coupled to each of the blade grips; a swashplate; and a
pitch link coupled to each of the pitch horns and the swashplate,
each of the pitch links being located closer to the mast axis than
an outermost surface of the yoke.
Description
BACKGROUND
When a helicopter is flying horizontally, or hovering in the wind,
differing relative wind speeds cause the rotating blades to
experience differing horizontal forces throughout each rotation.
For example, during forward flight, when the blade is advancing it
is encountering a larger relative air speed than when the blade is
retreating. Accordingly, each blade experiences large and varying
moments in the leading and lagging directions. Rather than rigidly
attaching blades to a yoke and forcing the yoke to absorb the large
varying moments, the blades may be attached to the yoke via a
lead-lag hinge which has an axis of rotation substantially parallel
to the mast axis. In order to prevent the blades from rotating too
far back and forth about the lead-lag hinge, and to prevent the
back and forth movement from matching the resonant frequency of the
drive system, dampers may be attached to the blades to provide a
resistive force.
The blades also experience large forces in a direction parallel to
the lead-lag hinge axis. In order to allow some movement in this
direction, a flap hinge may be utilized. The flap hinge attaches
the blades to the yoke about an axis perpendicular to the lead-lag
hinge axis.
In addition to the optional lead-lag and flap hinges, the blades
must be able to collectively and cyclically alter their pitch to
enable vertical and horizontal movement of the helicopter.
Therefore, each blade must be hinged about a pitch change axis that
is generally perpendicular to both the lead-lag hinge and flap
hinge axes.
The dampers may be coupled between the blades and the yoke or they
may be coupled between adjacent blades, known as blade-to-blade
dampers. Blade-to-blade dampers have generally been attached
proximate the trailing end of one blade grip and to the leading end
of the adjacent blade grip. As such, the attachment points of the
dampers are laterally offset from the pitch change axis. When the
blades are rotated away from horizontal, any resistive force
applied by the damper to the blade causes a rotational moment about
the pitch change axis. This moment must be resisted by the flight
control system in order to maintain the desired blade pitch. As the
blade rotates about the pitch change axis, the effective length of
the lever arm changes, and therefore, so does the moment. This is
further complicated by the constantly changing resistive force
which also modifies the magnitude of the moment. These constantly
changing moments unnecessarily complicate the dynamic analysis
required to effectively design and program the flight control
system.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is an oblique view of an aircraft comprising a rotor
assembly according to this disclosure.
FIG. 2 is an oblique view of a portion of the rotor assembly of
FIG. 1.
FIG. 3 is a top view of a portion of the rotor assembly of FIG.
1.
FIG. 4 is a top view of a portion of the rotor assembly of FIG.
1.
FIG. 5 is a side view of a portion of the rotor assembly of FIG.
1.
FIG. 6 is an oblique view of a portion of the rotor hub assembly of
FIG. 1.
FIG. 7 is an oblique view of a portion of the rotor hub assembly of
FIG. 1.
FIG. 8 is a top view of a portion of another rotor assembly
according to this disclosure.
FIG. 9 is a top cross-sectional view of a portion of the rotor
assembly of FIG. 9.
DETAILED DESCRIPTION
In this disclosure, reference may be made to the spatial
relationships between various components and to the spatial
orientation of various aspects of components as the devices are
depicted in the attached drawings. However, as will be recognized
by those skilled in the art after a complete reading of this
disclosure, the devices, members, apparatuses, etc. described
herein may be positioned in any desired orientation. Thus, the use
of terms such as "above," "below," "upper," "lower," or other like
terms to describe a spatial relationship between various components
or to describe the spatial orientation of aspects of such
components should be understood to describe a relative relationship
between the components or a spatial orientation of aspects of such
components, respectively, as the device described herein may be
oriented in any desired direction. In addition, the use of the term
"coupled" throughout this disclosure may mean directly or
indirectly connected, moreover, "coupled" may also mean permanently
or removably connected, unless otherwise stated.
This disclosure provides a novel rotor hub assembly that simplifies
the dynamic analysis required to design and program a flight
control system. This is accomplished with a rotor hub assembly that
utilizes dampers between adjacent blades to maintain in-plane
oscillations below, or above, 1/rev, i.e., below, or above, the
resonant frequency of the drive system. The dampers have attachment
points that are coincident to the pitch change axes of the blades.
In addition, the rotor hub assembly may utilize a single
axisymmetric elastomeric spherical bearing for each blade to serve
as the lead-lag, flap, and pitch hinges.
FIG. 1 illustrates an aircraft 100 comprising a main rotor assembly
102 according to this disclosure. Aircraft 100 comprises a fuselage
104 and rotor assembly 102 with a plurality of rotor blades 106.
Rotor assembly 102 is driven in rotation about a mast axis 108 by
torque provided by a powerplant housed within fuselage 104. Though
aircraft 100 is shown as a helicopter having a single main rotor,
rotor assembly 102 can alternatively be used on other types of
aircraft, such as, but not limited to, helicopters having more than
one main rotor or on tiltrotor aircraft. Also, rotor assembly 102
is shown as a main rotor for providing vertical lift and having
collective and cyclic control, though rotor assembly 102 may
alternatively be configured to provide longitudinal or lateral
thrust, such as in a helicopter tail rotor or airplane
propeller.
FIGS. 2 through 7 illustrate rotor assembly 102, various components
being removed for ease of viewing. A yoke 110 is coupled to a mast
112 for rotation with mast 112 about mast axis 108. Yoke 110 has a
honeycomb configuration in the embodiment shown, though in other
embodiments, yoke 110 may have a different configuration, such as a
central portion with radially extending arms. Yoke 110 is
preferably formed from a composite material, such as carbon fiber,
though yoke 110 may be formed from any appropriate material. In the
embodiment shown, yoke 110 is configured for use with six rotor
blades 106, though yoke 110 may be configured for use with any
appropriate number of blades. As shown in FIG. 2, yoke 110 may be
enclosed by an aerodynamic protective cover 114. Cover 114 may
include a top portion 116 and a bottom portion 118.
Yoke 110 has six bearing pockets 120, one bearing pocket 120
corresponding to each rotor blade 106. Each bearing pocket 120
carries a bearing 122, wherein bearing 122 may be an axisymmetric
elastomeric spherical bearing as disclosed in and described in U.S.
patent application Ser. No. 15/713,277 filed on Sep. 22, 2017, the
entirety of which is incorporated herein by reference. Each bearing
122 is spaced a radial distance from mast axis 108 and transfers
centrifugal force from the associated rotor blade 106 to yoke 110.
Each bearing 122 forms a lead-lag hinge to allow for limited
rotation of associated rotor blade 106 relative to yoke 110 in
in-plane lead and lag directions about a lead-lag axis, as
indicated by arrows 124 and 126, respectively. The lead-lag axis is
substantially parallel to mast axis 108 and passes through a center
point of each bearing 122. Bearing 122 also forms a flap hinge that
allows for limited rotation in out-of-plane flapping directions
about a flap axis, as indicated by arrows 128 and 130. The flap
axis is substantially perpendicular to the lead-lag axis and also
passes through the center point of bearing 122. Each bearing 122
also forms a pitch change hinge that allows for limited rotation
about a pitch change axis 132. Pitch change axis 132 is
substantially perpendicular to the lead-lag axis and the flap axis
and also passes through the center point of bearing 122. While each
rotor blade 106 can lead and lag about the associated bearing 122,
during operation the centrifugal force tends to force each rotor
blade 106 toward a centered, neutral position. It is from this
neutral position that each rotor blade 106 can lead, by rotating
forward (in the direction of rotation about mast axis 108,
indicated by arrow 124) in-plane relative to yoke 110, or lag, by
rotating rearward (indicated by arrow 126) in-plane relative to
yoke 110.
A blade grip 134 couples each rotor blade 106 to associated bearing
122, each blade grip 134 including an upper plate 136, a lower
plate 138, an inner portion 140, and a central portion 142. Inner
portion 140 and central portion 142 connect upper and lower plates
136, 138. As shown in the illustrated embodiment, inner portion 140
is a separate component that is coupled to upper and lower plates
136, 138, while central portion 142, upper plate 136, and lower
plate 138 comprise a unitary structure. Alternatively, inner
portion 140 and central portion 142 may be separate components that
are coupled to upper and lower plates 136, 138. Each blade grip 134
is connected to a proximal end 144 of a rotor blade 106 with
fasteners 146, thereby allowing loads from each rotor blade 106 to
be transferred through blade grip 134 and bearing 122 to yoke 110.
Fasteners 146 are inserted through blade attachment openings 148
extending through upper and lower plates 136, 138. Central portion
142 may include an aperture 150 extending therethrough. Proximal
end 144 of rotor blade 106 may cooperatively engage central portion
142 and/or aperture 150 to provide additional rigidity between
rotor blade 106 and blade grip 134.
A pitch horn 152 is coupled to each blade grip 134, allowing for
actuation by a pitch link 154 coupled between pitch horn 152 and a
swashplate 156 of a flight control system for causing rotation of
blade grip 134 and rotor blade 106 together about pitch change axis
132 for cyclic and collective control of rotor blades 106. Pitch
links 154 are oriented generally parallel to mast axis 108 and may
be located closer to mast axis 108 than the outermost portion of
yoke 110. Alternatively, pitch links 154 may be closer to mast axis
108 than the outermost portion of bearing pockets 120. Such a
configuration allows for a more compact, lightweight, aerodynamic
rotor assembly. Though not shown, a droop stop limits droop of each
rotor blade 106 and blade grip 134 assembly toward fuselage 104
when rotor assembly 102 is slowly rotating about mast axis 108 or
at rest.
Each rotor blade 106 is coupled to each adjacent rotor blade 106 by
a damper assembly 158, and each damper assembly 158 provides a
resistive force and cooperates with each adjacent damper assembly
158 to prevent large oscillations in lead-lag directions 124, 126,
and to maintain the frequency of in-plane oscillations below, or
above, 1/rev, i.e., below, or above, the resonant frequency of the
drive system. Damper assemblies 158 may be simple mono-tube
dampers, twin-tube dampers, hysteresis dampers, dry or wet friction
dampers, or magnetorheological dampers, wherein a magnetic field
may continuously modify the fluid viscosity, and thereby modifying
the damping properties. Damper assemblies 158 may provide
adjustable or fixed, as well as, linear or nonlinear resistance. A
connector, such as a rod end bearing 160, is installed at each end
of damper assembly 158. Rod end bearing 160 includes a ball 162
with a hole 164 extending therethrough. Ball 162 is housed within a
race 166. Rod end bearing 160 may also include a self-lubricating
liner between ball 162 and race 166 or it may include a zerk
fitting for the introduction of lubrication between ball 162 and
race 166.
To provide for coupling of damper assemblies 158 to blade grips
134, a first shaft 168, located adjacent to yoke 110, and a second
shaft 170, located adjacent to rotor blade 106, are rigidly coupled
to each blade grip 134 such that first and second shafts 168, 170
intersect pitch change axis 132. First shaft 168 and second shaft
170 are both sized for insertion through a respective hole 164 of
ball 162. Each ball 162 is coupled to either first shaft 168 or
second shaft 170 at the intersection of the respective shaft with
pitch change axis 132. When assembled, each damper assembly 158 can
be rotated a limited amount relative to each blade grip 134,
allowing for rotor blades 106 to rotate about pitch change axis 132
without materially affecting movement in lead and lag directions
124, 126 relative to each other and to yoke 110. The resistive
force of each damper assembly 158 is transferred to each blade grip
134 through associated rod end bearing 160, into first shaft 168 or
second shaft 170, and into adjacent blade grip 134 to resist
relative motion between blade grips 134 and their associated rotor
blades 106. Because rod end bearings 160 are coupled directly to
pitch change axis 132, the length of the lever arm between the
resistive force and pitch change axis 132 is zero. Therefore,
attachment directly to pitch change axis 132 effectively eliminates
any rotational moments that may be caused by the transmission of
force from damper assembly 158 to blade grip 134. If damper
assemblies 158 were coupled a distance away from pitch change axis
132, the forces applied by damper assemblies 158 would induce
rotation of rotor blade 106 about pitch change axis 132. Attachment
directly to pitch change axis 132 eliminates rotation, and
therefore, greatly simplifies the dynamic calculations required to
design and program the flight control system. It should be
understood that the attachment points of rod end bearings 160 need
not be directly on pitch change axis 132, as long as the attachment
points are close enough to pitch change axis 132 that the actual
lever arm is small enough that the moment created by forces from
damper assembly 158 are negligible when performing the required
dynamic analysis.
The configuration of rotor assembly 102 allows rotor blades 106 to
"pinwheel" relative to yoke 110, in which all rotor blades 106
rotate in the same lead or lag direction 124, 126 relative to yoke
110, and this may especially occur in lag direction 126 during
initial rotation about mast axis 108 of rotor assembly 102 from
rest. As the centrifugal force on rotor blades 106 builds with
their increased angular velocity, rotor blades 106 will rotate
forward in the lead direction 124 to their angular neutral position
relative to yoke 110. When damper assemblies 158 are configured as
shown in FIGS. 1-7, with a first rod end bearing 160 attached to
first shaft 168 proximate yoke 110 on a leading blade grip 134 and
a second rod end bearing 160 attached to a second shaft 170
proximate rotor blade 106 on a trailing blade grip 134, damper
assemblies 158 will provide resistive force to the pinwheeling
rotor blades 106. This occurs because second shafts 170 are further
away from the lead-lag axis, and therefore, second shafts 170
translate a larger distance from neutral when rotor blades 106
rotate in-plane than do first shafts 168, causing elongation of
damper assemblies 158 and the application of a resistive force. The
magnitude of the distance between first shaft 168 and second shaft
170 affects the amount of damping force applied during pinwheeling
of blades 106. Optionally, the distance between first and second
shafts 168, 170 is greater than or equal to the distance between
first shaft 168 and an outermost surface of yoke 110. The pinwheel
damping provided by staggered damper assemblies 158 eliminates the
need to include a filter in the full authority digital engine
control (FADEC) to prevent the extremely low in-plane frequency
common during pinwheeling from interfering with the engine control
frequency.
Referring to FIGS. 8 and 9, a rotor assembly 202 is shown. Rotor
assembly 202 is similar to rotor assembly 102 except that the
attachment points of the damper assemblies are not staggered. Rotor
assembly 202 includes a yoke 210 configured for use with rotor
blades 206. Each rotor blade 206 is coupled to each adjacent rotor
blade 206 by a damper assembly 258. Adjacent damper assemblies 258
include a concentric rod end bearing 260, which connects the ends
of two adjacent damper assemblies 258. Concentric rod end bearing
260 includes a ball 262 with a hole 264 extending therethrough.
Ball 262 is housed within an inner race 266, and inner race 266 is
housed within an outer race 268. Inner race 266 has a spherical
outer surface configured to freely rotate against a spherical inner
surface of outer race 268. Concentric rod end bearing 260 may also
include self-lubricating liners between ball 262 and inner race
266, as well as between inner race 266 and outer race 268.
To provide for coupling of damper assemblies 258 to blade grips
234, a single shaft 270 is rigidly coupled to each blade grip 234
such that shaft 270 intersects a pitch change axis 232. Shaft 270
is sized for insertion through hole 264 of ball 262. Each ball 262
is coupled to shaft 270 at the intersection of pitch change axis
232. When assembled, each damper assembly 258 can be rotated a
limited amount relative to each blade grip 234, allowing for rotor
blades 206 to rotate about pitch change axis 232 without materially
affecting movement in lead and lag directions 224, 226 relative to
each other and to yoke 210. The resistive force of each damper
assembly 258 is transferred to each blade grip 234 through
associated concentric rod end bearing 260, into shaft 270, and into
adjacent blade grip 234 to resist relative motion between blade
grips 234 and their associated rotor blades 206. Because concentric
rod end bearings 260 are coupled directly to pitch change axis 232,
the length of the lever arm between the resistive force and pitch
change axis 232 is zero. Therefore, attachment directly to pitch
change axis 232 effectively eliminates any rotational moments that
may be caused by the transmission of force from damper assembly 258
to blade grip 234. Attachment directly to pitch change axis 232
eliminates the rotational moment, and therefore, greatly simplifies
the dynamic calculations required to design and program the flight
control system.
At least one embodiment is disclosed, and variations, combinations,
and/or modifications of the embodiment(s) and/or features of the
embodiment(s) made by a person having ordinary skill in the art are
within the scope of the disclosure. Alternative embodiments that
result from combining, integrating, and/or omitting features of the
embodiment(s) are also within the scope of the disclosure. Where
numerical ranges or limitations are expressly stated, such express
ranges or limitations should be understood to include iterative
ranges or limitations of like magnitude falling within the
expressly stated ranges or limitations (e.g., from about 1 to about
10 includes, 2, 3, 4, etc.; greater than 0.10 includes 0.11, 0.12,
0.13, etc.). For example, whenever a numerical range with a lower
limit, R.sub.1, and an upper limit, R.sub.u, is disclosed, any
number falling within the range is specifically disclosed. In
particular, the following numbers within the range are specifically
disclosed: R=R.sub.1+k*(R.sub.u-R.sub.1), wherein k is a variable
ranging from 1 percent to 100 percent with a 1 percent increment,
i.e., k is 1 percent, 2 percent, 3 percent, 4 percent, 5 percent, .
. . 50 percent, 51 percent, 52 percent, . . . , 95 percent, 96
percent, 95 percent, 98 percent, 99 percent, or 100 percent.
Moreover, any numerical range defined by two R numbers as defined
in the above is also specifically disclosed. Use of the term
"optionally" with respect to any element of a claim means that the
element is required, or alternatively, the element is not required,
both alternatives being within the scope of the claim. Use of
broader terms such as comprises, includes, and having should be
understood to provide support for narrower terms such as consisting
of, consisting essentially of, and comprised substantially of.
Accordingly, the scope of protection is not limited by the
description set out above but is defined by the claims that follow,
that scope including all equivalents of the subject matter of the
claims. Each and every claim is incorporated as further disclosure
into the specification and the claims are embodiment(s) of the
present invention. Also, the phrases "at least one of A, B, and C"
and "A and/or B and/or C" should each be interpreted to include
only A, only B, only C, or any combination of A, B, and C.
* * * * *